Summary The flight departed Abbotsford International Airport, British Columbia, on an instrument flight rules flight plan at 1331 Pacific daylight time to conduct flight tests in the vicinity of Tofino, British Columbia. There were two pilots and two contract flight test engineers on board. A certification flight testing program was being conducted by Cascade Aerospace Inc. with the Bombardier DHC-8-402 (registration C-FBAM, serial number4040) modified for air tanker operations. The flight test program included a series of increasing amplitude push-over/pull-up manoeuvres to evaluate stick force perg (gravity) and to demonstrate satisfactory handling characteristics during a push-over to 0g with a water/retardant load on board. At about 1532, during the third and final push-over, both the No.1 and No.2 engine low oil pressure annunciators illuminated. Then, uncommanded, the speed of both propellers reduced by about 300rpm, and the engine torques increased proportionately. A recovery to a normal attitude was completed during which both engine low oil pressure annunciators extinguished. The No.1 propeller recovered to the selected speed of 1020rpm. However, the No.2 propeller went into an overspeed condition and was governed at 1060rpm by the overspeed governor. The No.2 propeller was then intentionally feathered, the engine was shut down, and the water/retardant load was jettisoned. The No.2 engine was subsequently re-started. However, the propeller could not be unfeathered, so the engine was again shut down. The flight returned to Abbotsford International Airport with one engine inoperative and landed at about 1640. Air traffic control was advised of the condition, but an emergency was not declared. Ce rapport est galement disponible en franais. Other Factual Information Meteorological Information The weather conditions for Tofino at the time of the incident were as follows: wind 270true(T) at 11knots gusting to 17knots; visibility 15statute miles; a few clouds at 2400feet above sea level, scattered cloud at 5000 feet, broken cloud at 15000feet, temperature 15C, dew point 10C. Operations Cascade Aerospace Inc. provides engineering services for custom product design, manufacturing, and certification. The company is an approved Design Approval Organization (DAO) as well as an Approved Manufacturing Organization (AMO). The Q400Airtanker conversion of the initial Bombardier DHC-8-400 design was undergoing supplemental type certificate (STC) certification flight testing for the fire-fighting role at the time of the incident. The pilot crew was employed by the Conair Group Inc. Both pilots had completed the Dash8 Q400initial pilot course at the FlightSafety International Toronto Learning Centre in November2004. Aircraft The Bombardier DHC-8-400 is a current production regional airliner in the 70seat market. It is powered by two Pratt& Whitney Canada (P) PW150A turboprop engines and is operated by a two-pilot crew. The Cascade Aerospace Inc. Q400Airtanker conversion included the installation of a 10000litre retardant delivery system. The aircraft departed Abbotsford International Airport at a gross weight of 68200pounds and the load was configured at the most aft centre of gravity limit for the test flight. Part of the STC modifications included an increase in the aircraft gross take-off weight to 68200pounds. The aircraft was being operated under the authority of a Transport Canada (TC) Flight Permit- Experimental. This was the eighth test flight in the development of an STCfor an aerial tanker conversion. Subsequent to the flight, a maintenance interrogation of the recorded engine monitoring unit (EMU) data did not identify any component faults. Flight Recorders The aircraft was equipped with a Honeywell solid-state cockpit voice recorder (CVR) and an Allied Signal solid-state flight data recorder (FDR). Information was recovered from the FDR but not from the CVR. The pertinent information from the CVR was overwritten while external electrical power was applied to the aircraft following the incident flight. Powerplants The aircraft was equipped with two P model PW150A engines. The PW150A consists of a free-turbine turbo machine (TM) module driving a Dowty Aerospace model R408 six-bladed propeller through a two-stage reduction gearbox (RG) module. The TM module includes the low-pressure (LP) compressor and its LPturbine, the single-stage centrifugal high-pressure (HP) compressor and its HP turbine, and the two-stage power turbine (PT) and its PT shaft, which drives the RGmodule. The three rotating assemblies are not connected together and rotate at different speeds and in opposite directions. An accessory gearbox (AGB) is driven by the HP turbine shaft through a tower shaft and angle gearbox. The AGB drives the fuel metering unit, which incorporates the fuel pump, the starter/generator and other accessories. Engine Pitch Angle Limit A series of flight test cards was produced during the development of the flight test program. A flight test card outlined the objectives, procedures, and safety concerns of each flight. The Cascade Aerospace Inc. flight test engineering group referenced flight test guidance material produced by TC, the United States Federal Aviation Administration (FAA) and the manufacturer's aircraft flight manual (AFM). The AFM was the primary document used to determine aircraft operating limitations. Section2.2.5of the AFM, Manoeuvring Limit Load Factors, indicates that the load factor limits for the aircraft in the clean configuration are +2.5g to -1.0g. The section further states that these figures limit the permissible bank angle in turns and limit the severity of pull-up and push-over manoeuvres. Cascade Aerospace Inc. self-imposed another condition on the TC flight permit that limited the minimum manoeuvre limit load factor to 0g. Flight test card No. 8 did not contain any pitch angle limitations. Data recovered from the FDR indicated that, during the manoeuvres, pitch-up angles of about 43and 46were achieved in the second and third pull-ups respectively, and a minimum load factor of -0.08g was recorded during the last push-over. The FDR and the captain's attitude indicator normally receive pitch information from the same-side attitude and heading reference unit. The engine manufacturer's Engine Installation Manual (EIM) (a proprietary document provided to the aircraft manufacturer but not normally distributed to aircraft owners/operators) contained static engine operating limitations regarding the demonstrated pitch and roll angles and related time frames at which satisfactory engine lubrication could be lost. This manual was not offered to, nor requested by, the Cascade Aerospace Inc. flight test engineering group. Although the aircraft did not exceed the limit load factor prescribed in the AFM, the engine static pitch angle limitation was, unbeknownst to the flight test crew, exceeded during the flight test manoeuvre. Although the physical pitch-up angle reached as high as 46, when acceleration and deceleration forces are taken into account, the net effect on the engine was equivalent to a steady-state pitch-up angle of about 8. The pitch-up limitation for normal steady-state operation is 25according to the EIM. Engine Oil Pressure/Propeller Relationship Each propeller and engine (combined RG module plus TM module) shares the same oil, so when both engines lost oil pressure, both propellers did as well. The loss of oil pressure resulted in each propeller counterweight mechanism and rotational forces driving the respective propeller in the coarse pitch direction and resulted in an underspeed condition (80percent Np[power turbine]). This event was not an auto-feather since the auto-feather system had been disarmed as per normal procedures. The FDR data show that, before the loss of oil pressures, the engines were operating at less than 50percent torque. At about the time the engine oil pressures were recovering and the low oil pressure annunciators were being extinguished, the No.2 power lever was advanced about 8(48to56) followed shortly by the advance of the No.1 power lever. During this period of two or three seconds, the No.2 engine torque increased to above 50percent, and the conditions (torque 50percent and propeller speed 80percent for more than one second) were fulfilled to activate the hardware-based automatic underspeed protection circuit (AUPC) for the No.2 propeller system. This circuit effectively disables the propeller electronic control (PEC) software control and continuously drives the propeller blade pitch at a controlled rate in the fine direction, and forces it to operate on the hydro-mechanical overspeed governor at 1060rpm. A similar AUPC activation and respective fault code was not recorded for the No.1 engine, indicating that the conditions required for the activation of the No.1 AUPC were not met during the transient. In the case of the No.1 engine, once oil pressure was restored, the propeller simply resumed normal operation. Propeller Unfeathering After confirming that the No. 1 engine/propeller system was operating normally, the crew complied with the checklist in the AFM by manually feathering the No.2 propeller and shutting down the No.2 engine. Following an in-flight maintenance consultation, the No.2 engine was subsequently re-started; the propeller could not be unfeathered so the engine was again shut down. In accordance with the design of the system, the propeller was inhibited from unfeathering following the AUPC activation. The philosophy of the aircraft manufacturer is that certain propeller system fault conditions are automatically accommodated by the control system, feathering the propeller until the fault conditions are corrected by appropriate maintenance action. The AUPC triggers this accommodation mode and prevents propeller unfeathering since its activation is equivalent to a loss of blade pitch control by the PEC software. Aircraft and Maintenance Records Aircraft records indicate that the left engine, serial number PCE-FA0020, had about 1952 flight hours of total time since new (TTSN) and the right engine, serial number PCE-FA0015, had about 1991flight hours of TTSN. Service Bulletin (SB)35038, Revision3, had been incorporated into both engines and was applicable to all production PW150A engines. This SBrequired replacement of the original engine oil pressure relief valve (PRV) with a modified PRV. The modification was in response to an oil smell in the passenger cabin due to oil leakage into the bleed-air stream while engine bearing seals were unpressurized during the start-up and shutdown sequence. The function of the modification was to prevent excessive oil supply to the engine bearing cavities, which could cause oil flooding during engine start-up and shutdown by relieving residual oil pressure more quickly. Research A search of the TC service difficulty reporting database did not produce any similar occurrences in Canada or the United States. At this time, the only civilian aircraft equipped with the PW150A engine is the DHC-8-400 model. During this investigation, it was noted that the outcome in this occurrence was different from the outcome of previous 0g certification flight test demonstrations conducted by Bombardier in1999. Following this incident, P conducted a series of tests on an experimental engine of the same model to determine whether the PRV modification (SB35038, Revision3) affected the propeller operation and how this sequence of events compared to the original certification flight test demonstrations with the originalPRV. The original PRV configuration was installed in the engine used in the P propeller test stand. During those demonstrations, the power plant and propeller continued to perform as it did during the original certification tests throughout the artificially induced temporary low engine oil pressure condition. This testing artificially induced the loss of oil pressure by introducing pressurized air at the oil pump inlet until the main oil pressure dropped to 10psi, and, therefore, any effects of flight under conditions of less than 1g were not present. With the original PRV configuration, tested at two different power levels, there was no change in propeller operation throughout the loss of oil pressure and recovery cycle. With the modified PRV configuration, there were fluctuations in propeller operation, and the engine main oil pressure dropped to below 10psi and took longer to recover. The propeller operation recovered about four seconds after the main and reduction gear oil pressure recovered. The testing confirmed that the modified PRV does cause the propeller blade angle to move toward the feather position under conditions of low engine oil pressure. The propeller does recover to the prior operating state once oil pressure is restored. However, a loss of oil pressure for a duration longer than 3or 4seconds may result in activation of the AUPC.